This invention relates generally to fatigue-resistant and damage-tolerant components and methods of producing such components.
Various metallic, ceramic, and composite components, such as gas turbine engine fan and compressor blades, are susceptible to cracking from fatigue and damage (e.g. from foreign object impacts). This damage reduces the life of the part, requiring repair or replacement.
It is known to protect components from crack propagation by inducing residual compressive stresses therein. Methods of imparting these stresses include shot peening, laser shock peening (LSP), pinch peening, and low plasticity burnishing (LPB). These methods are typically employed by applying a “patch” of residual compressive stresses over an area to be protected from crack propagation, for example a leading edge of a gas turbine engine compressor blade.
During a burnishing operation, the depth of the compressive residual stress layer can be controlled with process parameters. It is known to control those parameters to transition from high stress areas to low stress areas to prevent a high gradient from compressive to tensile stress fields (this technique is known as “feathering”). However, through the rest of the process, the parameters are held constant, even when processing non-uniform cross-sections (triangular cross-sections, for example). This can result in areas of tensile stresses between layers of compressive residual stress, along with areas where the compressive residual stresses are higher than the intended range.
FIG. 1 illustrates a generic metallic component 10 with a surface 12. A burnishing element 14 is pressed against the surface under substantial pressure and translated along a selected path. In this example the burnishing element 14 is a sphere, but cylindrical rollers are also used. Typically a pressurized fluid is used to force the burnishing element 14 onto the surface 12 of the component 10. Mechanically loaded tools are also used. Appropriate equipment, of a known type, typically CNC controlled, is provided to load the burnishing element 14, and to move it along the desired path. The pressing force used during burnishing is such that it induces plastic strain and a region of residual compressive stresses 16 within the component 10 near a burnished area 18. A region of residual tensile stresses 17 exists around the border of the region 16.
FIG. 2 illustrates an exemplary gas turbine engine compressor blade 20. This component is used merely as an example of a part to which both prior art methods and the present invention may be applied. the present invention is equally applicable to other types of components susceptible to cracking from fatigue or damage, such as compressor stator vanes, fan blades, turbine blades, shafts and rotors, stationary frames, actuator hardware and the like. Such components may be made from metal alloys, ceramics, or composite materials (e.g. carbon fiber composites). The compressor blade 20 includes an airfoil 22, a platform 24, and a shank 26. In this particular example the shank 26 includes a dovetail 28 for being received in a slot of a rotating disk (not shown). The airfoil 22 has a leading edge 30, a trailing edge 32, a tip 34, a root 36, a pressure side 38, and a suction side 40 opposite the pressure side 38. A burnishing tool 42 carrying a burnishing element 14 is shown tracing out a selected burnishing path “P” along the surface of the airfoil 22. In this example, the path “P” includes a plurality of linear segments 23 arranged in a series of S-turns. The path has a footprint with a width “W” determined by the width of the burnishing element 14 and the applied pressure. The linear segments 23 are separated by an step-over distance “S”. In cases where the step-over distance S is less than the width W, overlap of the segments 23 will occur. In most applications, there will be substantial overlap to achieve adequate coverage and desired stress profiles.
FIGS. 3A and 3B illustrate a prior art burnishing treatment being applied to edge 32 of the airfoil 22. FIG. 3A shows the treatment being applied to the pressure side 38 by a single burnishing element 14, while the airfoil is supported by a block 44. In this case, a constant applied pressure in the normal direction “f” is selected to generate a region 46 of residual compressive stress which has depth “d” defined as a distance from the surface of the pressure side 38, expressed as a fraction of the total thickness of the airfoil 22 at the point of measurement. The burnishing element 14 is moved from left to right. The depth d will decrease substantially as the burnishing element 14 traverses the thicker portion of the airfoil 22 distal from the trailing edge 32. The result is that the interior boundary 48 of the region 46 is not parallel to a mid-chord plane M of the airfoil 22. Under these circumstances, the depth d will vary significantly from a desired magnitude at opposite axial ends of the region 46, regardless of which end is used as the basis for setting the applied pressure.
FIG. 3B illustrates the prior art burnishing treatment being applied to both the pressure side 38 and the suction side 40 of the airfoil 22 by opposed burnishing elements 14 and 14′. In this case, the applied pressure in the normal directions, denoted f and f′, are selected to generate regions 50 and 52 of residual compressive stress which have depths d and d′ measured from the surface of the pressure side 38 and suction side 40, respectively, and expressed as a fraction of the total thickness of the airfoil 22 at the point of measurement. The depths d and d′ are typically chosen to generate through-thickness residual compressive stress near the trailing edge 32. However, as shown, the depths d and d′ will decrease substantially as the burnishing elements 14 and 14′ traverse the thicker portion of the airfoil 22 distal from the trailing edge 32. The result is that the interior boundaries 54 and 56 of the regions 52 and 54 are not parallel to a midplane M of the airfoil 22. If the pressures f and f are just enough that through-thickness residual compressive stress is produced near the trailing edge 32, this results in an internal region 58 of residual tensile stress at thicker portions of the airfoil 22. It is possible to select the pressures f and f′ so that the regions 50 and 52 merge to produce through-thickness residual compressive stress, even at the thickest portion of the treated area. However, this would result in excessive compressive stress levels near the trailing edge 32, because of overlap of the regions 50 and 52. It could also damage the airfoil 22 and result in undesired deformation.
In light of the above shortcomings of the prior art, there is a need for a method of producing uniform through-thickness residual compressive stresses in components of variable thickness.